This invention relates to gas turbine engines and more particularly to axial flow compressors for such engines.
An axial flow compressor generally comprises one or more rotor assemblies that carry blades of aerofoil section, the rotor assemblies are carried within a casing within which are located stator blades. The compressor is a multi-stage unit, as the amount of work done (pressure increase) by each stage is small; a stage consists of a row of rotating blades followed by a row of stator blades. The reason for the small pressure increase across each stage is that the rate of diffusion and the deflection angle of the blades must be limited if losses due to air breakaway of the blades and subsequent blade stall are to be avoided.
The condition known as stall, or surge, occurs when the smooth flow of air through the compressor is disturbed. Although the two terms "stall" and "surge" are often used synonymously, there is a difference which is mainly a matter of degree. A stall may affect only one stage or even group of stages, but a compressor surge generally refers to a complete flow breakdown through the compressor.
The value of airflow and pressure ratio at which a surge occurs is termed the "surge point". This point is a characteristic of each compressor speed, and a line which joins all the surge points, called the surge line (FIG. 7), defines the maximum stable airflow which can be obtained at any rotational speed. A compressor is designed to have a good safety margin (Region A) between the airflow and the pressure ratio at which it will normally be operated (the working line), and the airflow and pressure ratio at which a surge will occur.
For satisfactory operation of a compressor stage, it is well known that it, and also the adjacent stages of the blades, must be carefully matched as each stage possesses its own individual airflow characteristics. Thus it is extremely difficult to design a compressor to operate satisfactorily over a wide range of operating conditions such as an aircraft engine encounters.
Outside the design conditions, the gas flow around the blade tends to degenerate into a violent turbulence, and the smooth pattern of flow through the stage or stages is destroyed. The gas flow through the compressor usually deteriorates and becomes a rapidly rotating annulus of pressurized gas about the tips of one compressor blade stage or group of stages. If a complete breakdown of flow occurs through all the stages of the compressor such that all the stages of blades becomes "stalled", the compressor will "surge".
The transition from stall to surge can be so rapid as to be unnoticed, or on the other hand, a stall may be so weak as to produce only slight vibration or poor acceleration or deceleration characteristics. A more severe compressor stall is indicated by a rise in turbine gas temperature, and vibration or "coughing" of the compressor. A surge is evident by a bang of varying severity from the engine compressor and a rise in turbine gas temperature.
It is necessary to use a system of airflow control to ensure the efficient operation of an engine over a wide speed range and to maintain the safety margin referred to above. A well known method of control is described in British Patent 1,518,293 and consists of providing the compressor casing of such an engine with a circumferential row of slots inclined to the axis of rotation of the rotor blade row and disposed within its internal cylindrical surface adjacent to at least one blade row. The slots have an axial length substantially greater than that of the blade row, and terminate downstream of the blade row.